WO1999005407A1 - Agencement de deflecteur pour dispositif de suppression d'infrarouges - Google Patents

Agencement de deflecteur pour dispositif de suppression d'infrarouges Download PDF

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Publication number
WO1999005407A1
WO1999005407A1 PCT/US1998/014212 US9814212W WO9905407A1 WO 1999005407 A1 WO1999005407 A1 WO 1999005407A1 US 9814212 W US9814212 W US 9814212W WO 9905407 A1 WO9905407 A1 WO 9905407A1
Authority
WO
WIPO (PCT)
Prior art keywords
turning vane
flow
turning
defines
gaseous fluid
Prior art date
Application number
PCT/US1998/014212
Other languages
English (en)
Inventor
Mark H. Hammond
James H. Riffanacht
Original Assignee
Sikorsky Aircraft Corporation
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Sikorsky Aircraft Corporation filed Critical Sikorsky Aircraft Corporation
Priority to DE69805800T priority Critical patent/DE69805800T2/de
Priority to AU82957/98A priority patent/AU8295798A/en
Priority to JP2000504365A priority patent/JP2001511495A/ja
Priority to EP98933276A priority patent/EP0998628B1/fr
Publication of WO1999005407A1 publication Critical patent/WO1999005407A1/fr

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/46—Nozzles having means for adding air to the jet or for augmenting the mixing region between the jet and the ambient air, e.g. for silencing
    • F02K1/48—Corrugated nozzles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
    • B64D33/04—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of exhaust outlets or jet pipes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/78—Other construction of jet pipes
    • F02K1/82—Jet pipe walls, e.g. liners
    • F02K1/822—Heat insulating structures or liners, cooling arrangements, e.g. post combustion liners; Infrared radiation suppressors
    • F02K1/825—Infrared radiation suppressors
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
    • B64D33/04—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of exhaust outlets or jet pipes
    • B64D2033/045—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of exhaust outlets or jet pipes comprising infrared suppressors
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This invention is directed to InfraRed (IR) Suppressors for engines, and more particularly, to a turning vane arrangement therefor operative to suppress the infrared radiation emitted by an engine.
  • IR InfraRed
  • the exhaust plume of gas turbine engines is a source of high infrared energy which may be used for targeting/tracking purposes by heat seeking missiles and/or various forms of infrared imaging systems, e.g.. Night Vision Goggles (NVGs).
  • heat seeking missiles obtain directional cues from the infrared energy, which directional cues are more precise depending upon the intensity, e.g.. temperature, and total energy emitted by the exhaust plume.
  • infrared imaging systems amplify the infrared energy by a factor of about 10,000 and may be used by hostile forces for early detection and/or targeting of surface-to-air/air-to-air missiles. Accordingly, it is highly desirable to reduce the infrared emissions (also referred to as the IR signature) of such engines to a practical minimum to avoid detection, acquisition and/or tracking by enemy threats/forces.
  • infrared suppression systems have been designed and fielded which effect mixing of low temperature ambient air with the high temperature engine exhaust to reduce the IR signature radiated therefrom.
  • Secondary objectives may include: (i) minimizing system or configuration complexity to reduce fabrication costs, (ii) minimizing the external aerodynamic drag produced by such IR suppressors, and/or (iii) suppressing the acoustic emissions emanating therefrom which may also be a source of detection.
  • Hurley U.S. Patent 4.018,046 discloses an InfraRed (IR) suppressor for reducing the infrared radiation emitted from the exhaust plume and/or hot metal parts of a gas turbine engine.
  • the IR suppressor includes an ejector vane assembly which is adapted for receiving a primary flow of engine exhaust and which is operative to entrain a first flow of cooling air into the primary exhaust flow.
  • the primary flow of gases diffused from and entrained by the ejector vane assembly are fed into an arcuate duct which, due to its spatial position relative to the ejector vane assembly, produces an annular space for entraining a second flow of cooling air.
  • the first and second flows are admixed with the primary exhaust flow to effect heat transfer therebetween and. consequently, to suppress the IR signature emitted from the exhaust plume.
  • the curved shape of the arcuate duct and the use of turning vanes disposed internally thereof serve to prevent direct line-of-sight viewing into the IR suppressor. That is. the shape and/or occluding characteristics of the turning vanes prevents viewing of radiant energy emitted by hot internal components such as from the ejector vane assembly or from the turbine section of the engine.
  • the ejector vane assembly described therein includes a plurality of V-shaped ejector vanes which entrain cooling air at a plurality of angularly spaced positions while additionally serving to introduce air toward the center or "core" of the engine exhaust. More specifically, each of the ejector vanes forms a V-shaped channel which is open to ambient air at one end thereof and is oriented to direct flow radially toward the core of the primary exhaust flow.
  • the ejector vane assembly furthermore, clusters the ejector vanes as close as practicable to the core of the priman, exhaust flow and. furthermore. includes a means for adjusting the flow area through the ejector vanes to ameliorate engine performance. That is. the adjustment means effects collapse of each ejector vane so as to increase the flow area and reduce the back-pressure on the engine.
  • the ejector vane assembly improves mixing, i.e.. reduces the IR signature, by directing cooling air into a core region of the engine exhaust
  • the ejector vane assembly requires a large number of individual components.
  • the multiplicity of components each being a potential source of repair and maintenance, reduces the overall reliability of the IR suppressor system.
  • the adjustment means attempts to reduce flow restriction, the anticipated power loss produced by the ejector vane assembly is on the order of between 3% to 6%.
  • power loss may be viewed as intolerable for certain aircraft and/or aircraft missions wherein maximum gross weight, flight speed, and range are critical design parameters.
  • the turning vanes serve to redirect the exhaust flow and prevent direct line-of-sight viewing internally of the arcuate duct, impingement of the high temperature exhaust gases on the surfaces thereof can cause the turning vanes to become a significant source of infrared radiation. That is, depending upon the temperature reduction effected by the first stage of the suppressor, i.e., the ejector vane assembly, the temperature of the turning vanes can significantly contribute to the IR signature of the suppressor.
  • Miller U.S. Patent 4,312,480 discloses a gas diffusion and radiation shielding apparatus for a turbine engine which provides multiple stages for entraining cooling air into a primary flow of high temperature engine exhaust, and a plurality of channels for segmenting the primary flow into smaller, more efficient, mixing streams.
  • each internal conduit includes an inlet disposed along the upper and lower edges of the channel and an outlet disposed along a leading edge which opposes the primary flow of exhaust gases.
  • the high-velocity low- pressure primary flow draws a "reverse flow" of cooling air through the conduit which is diffused via the leading edge outlet.
  • the cooling air reduces the surface temperature of the sidewalls and, consequently, the IR signature.
  • IR radiation emitted by the channels, and more specifically, the sidewalls is reduced via the reverse flow of cooling air, the reduction in IR signature is marginally beneficial, especially when compared to the cost of fabricating such internal conduits. Furthermore, by directing cooling air outwardly from the leading edge of the sidewall, back-pressure is produced which degrades engine performance.
  • Presz et al. U.S. Patents 4,835,961 and 4,830,315 describe nozzle configurations for mixing/pumping fluid. More specifically, the Presz '961 and '315 patents describe single and multi-stage mixer/ejector nozzles, respectively, each having a plurality of adjoined lobes formed at an outlet end thereof. The adjoined lobes define a primary flow trough for channeling a primary flow of high velocity/energy fluid, such as the exhaust of a gas turbine engine, and a secondary flow trough for channeling a secondary flow of low velocity/energy fluid such as ambient air.
  • the primary and secondary flow troughs are alternately disposed about the periphery of the nozzle such that thin sheets of high energy fluid flow from the trough outlets, transfer kinetic energy to the low energy fluid, and entrain the secondary flow into the primary flow.
  • the adjoined lobes produce axial vortices which rapidly admix the primary and secondary flows.
  • the adjoined lobes described in the Presz '961 and '315 patents are known to have utility on aircraft engines to both increase thrust and for suppressing the thermal energy radiated from the engine.
  • the nozzle configurations disclosed therein fail to fully mix the flows, particularly in the core region of the primary exhaust flow. That is, in all of the embodiments described therein, the lobes entrain the secondary cooling flow into a perimeter region of the flow such that a central core or band of high temperature exhaust remains as the flow exits the nozzle. While the core of high temperature flow will eventually admix downstream, it will be appreciated that the degree of IR suppression is compromised until such energy is diffused. Aside from the IR emissions radiated by the engine, yet another source of detection includes the observable "radar signature" of the aircraft.
  • the radar signature of an aircraft may be defined as the level of electromagnetic energy returned, i.e., reflected back, to a scanning radar when the aircraft is in the detection range or interrogation field of the radar.
  • the scanning radar may provide early detection so that appropriate countermeasures may be taken, or, if the radar signature is below the detection level of the scanning radar, may be entirely ineffective. Accordingly, for aircraft conducting covert operations, it is highly desirable to reduce the radar signature of the aircraft to obviate an enemy's ability to conduct countermeasures.
  • IR suppressor outlet which placement prevents radar line-of-sight viewing relative to an anticipated radar azimuth angle. For example, by placing the outlet upward, the aircraft fuselage may effectively occlude the suppressor outlet. Notwithstanding the method employed, it will be appreciated that the impact on radar signature is exacerbated on aircraft requiring multiple IR suppression apparatus.
  • IR InfraRed
  • a turning vane arrangement for an IR suppressor having a duct adapted for receiving a primary flow of engine exhaust.
  • the turning vane arrangement is operative to direct a pressurized secondary flow of low temperature gaseous fluid ⁇ into the primary flow of engine exhaust for reducing the IR energy thereof.
  • the turning vane arrangement includes at least one turning vane situated in the duct and is characterized by a trailing edge that defines an exhaust nozzle.
  • the exhaust nozzle is adapted for being disposed in fluid communication with a flow source which produces the pressurized secondary flow.
  • the exhaust nozzle may include a plurality of adjoined lobes for rapidly and thoroughly admixing the primary and pressured secondary flows.
  • Fig. 1 is a profile view of a helicopter illustrating an exemplary installation of an IR suppressor employing a turning vane arrangement according to the present invention
  • Fig. 2 is a partially broken-away perspective view of the IR suppressor for revealing the internal details of the turning vane arrangement and for illustrating the various fluid flow paths through and around the turning vanes arrangement;
  • Fig. 3 is an isolated perspective view of a single turning vane employed in the turning vane arrangement and the exhaust nozzle configuration employed along the trailing edge thereof;
  • Fig. 4 is an end view of the turning vane arrangement for defining the core region of the primary exhaust flow and the placement of the turning vane nozzles relative thereto;
  • Figs. 5a and 5b depict various alternate embodiments of the exhaust nozzle configuration useful in the turning vane arrangement according to the present invention.
  • exemplary embodiments of the present invention will be described in the context of a helicopter gas turbine engine, though, it will be appreciated that the teachings described herein are applicable to any engine/vehicle wherein the suppression of InfraRed (IR) energy is desired.
  • IR InfraRed
  • the invention will be described in the context of an IR suppressor having multiple ejector stages, though it will be appreciated that the teachings are applicable to single or multi-stage mixer/ejectors.
  • the described embodiment of the invention relates to integrating the exhaust flows of a turbo-shaft engine and a transmission oil cooler, the invention may be used in combination with a single engine, multiple engines or multiple exhaust flows, whether or not produced by an engine.
  • Fig. 1 shows a helicopter 10 having one or more gas turbine engines 12 for driving a main and tail rotor, 14 and 16, respectively.
  • An IR suppressor 20 embodying features of the present invention is disposed in combination with the aft end of each gas turbine engine 12 and is operative to suppress the InfraRed (IR) signature radiated from the high-temperature exhaust of the respective engine 12.
  • IR InfraRed
  • to suppress means that the IR signature emanating from the gas turbine engine 12 is reduced below some predetermined threshold value which is indicative of the acquisition, tracking and/or targeting capability of a particular IR threat.
  • the oil cooler 22 for cooling the lubricating oil circulated through a main helicopter transmission (not shown).
  • the oil cooler 22 includes a heat exchanger and fan (not shown) and exhausts a flow of pressurized air having a temperature of about 180 - 220 degrees Farenheit.
  • Such exhaust flow may be characterized as "low temperature" relative to the high temperature exhaust of the engine 12. which exhaust is about 1200 to 1400 degrees Farenheit.
  • a partially broken-away perspective view of the IR suppressor 20 is shown in combination with the aft end of the engine 12 and in position relative to the helicopter fuselage 26.
  • the IR suppressor 20 comprises a first stage exhaust nozzle 28 and a turning vane arrangement 30 according to the present invention situated downstream thereof.
  • the IR suppressor includes a shroud/duct 32 which envelopes the first stage exhaust nozzle 28 and which surrounds and supports the turning vanes 30.
  • the first stage exhaust nozzle 28 includes an inlet conduit 34 which is disposed in combination with the engine exhaust casing 36 and a plurality of adjoined lobes 38 which form an outlet end of the first stage exhaust nozzle 28.
  • the first stage exhaust nozzle 28 is adapted for receiving a primary flow PF of high temperature exhaust E HT from the engine 12 and diffuses the same through the adjoined lobes 38.
  • the effluent primary flow PF entrains a secondary flow SF of low temperature gaseous fluid, e.g., ambient air, through an annular flow passage 40 defined by the shroud/duct 32.
  • the fluid dynamics principle for entraining such secondary flow SF is well-known in the art and is often referred to as a "venturi effect".
  • the wave-shape configuration of the adjoined lobes 38 engenders the formation of axial vorticies which rapidly admixes the primary and secondary flows PF, SF to produce a primary flow PF of partially cooled exhaust E PC .
  • Such lobe configuration and mixing principle is known in the art and more fully described in Presz et al. U.S. Patent 4.835,961.
  • the turning vane arrangement 30 comprises at least one turning vane 50 having a trailing edge 50 TC which defines an exhaust nozzle 52.
  • the exhaust nozzle 52 is disposed in fluid communication with a flow source 54 which provides a pressurized secondary flow PSF of low temperature gaseous fluid.
  • a fluid communication means which comprises an internal conduit 56 disposed within the exhaust nozzle 52 and an external conduit 58 (Fig. 2) disposed between the flow source 54 and the internal conduit 56.
  • the flow source 54 is any low- temperature, fluid flow source capable of imparting momentum to the primary flow PF of engine exhaust.
  • the flow source 54 is the transmission oil cooler 22 (Fig. 1)
  • the exhaust nozzle 52 of each turning vane is adapted for being disposed in a "core region" of the shroud/duct 32.
  • the "core region” CR means the centermost 30% of the flow area, i.e., the area defined and bounded by the inner wall 32 w of the shroud/duct 32, at an axial location corresponding to the trailing edge 50- ⁇ of the turning vain 50.
  • the core region of the shroud/duct also corresponds to the core region of the primary exhaust flow PF which is typically the hottest portion of the primary flow PF.
  • the exhaust nozzle 52 is configured and situated along the trailing edge 50- ⁇ so as to span a length L [ which corresponds to its intersection of the core region CR, i.e., the turning vane intersection length. More specifically, the exhaust nozzle 52 spans a length L 52 which is equal to or greater than the intersection length L,.
  • the exhaust nozzle 52 may be a simple elongate slot
  • the exhaust nozzle 52 of the preferred embodiment defines a plurality of adjoined lobes 60. More preferably, the adjoined lobes 60 define a reflective lobe configuration 60 R wherein a "mirror-image" of the lobe configuration 60 R is provided on both sides of the exhaust nozzle. The import of such lobe configuration 60 R will be discussed in subsequent paragraphs.
  • the turning vane arrangement 30 provides a number of mechanisms, some known in the art and others which are the subject of this invention, to reduce the IR signature radiated by the engine 12.
  • the turning vane arrangement 30 serves to redirect the primary flow so as to prevent impingement thereof on other fuselage structure. Secondly, in conventional fashion, the turning vane arrangement 30 prevents direct line-of-sight viewing into the interior of the IR suppressor 20. Accordingly, radiant energy emitted by hot internal components such as the turbine section of the engine 12 is occluded from viewing by IR sensing devices/threats.
  • the turning vane arrangement 30 provides several mechanisms to rapidly and thoroughly mix/cool the primary exhaust flow PF. More specifically, the pressurized secondary flow PSF of low temperature gaseous fluid is caused to flow out the trailing edge exhaust nozzle 52 for admixture with the primary exhaust flow PF. More specifically, and referring to Fig. 3, the adjoined lobes 60 of the exhaust nozzle 52 cause the primary and secondary flows PF, PSF to penetrate each other so as to produce a plurality of axial vortices downstream of the exhaust nozzle 52.
  • the vortices engendered by the lobe configuration effect rapid and thorough admixture of the primary and secondary flows PF, PSF, such that the average temperature of the resultant cooled engine exhaust E c is about 1/2 of the temperature of the partially cooled exhaust E PC .
  • the exhaust nozzles 52 are situated in the core region CR of the primary exhaust flow. As such, the pressurized secondary flow PSF is introduced into the hottest portion of the primary exhaust flow PF so as to diminish the IR signature to a maximum practical extent.
  • the internal conduit is configured so as to occupy a substantial portion of the internal volume of the turning vane 50.
  • the wall thickness T of the turning vane 50 should be minimized to maximize the rate of heat transfer occurring between the exterior surfaces 50 s of the turning vane 50 and an interior surface 56 s of the internal conduit 56. It will be apparent to one of ordinary skill that an optimum wall thickness T may be determined based on conventional principles of heat transfer.
  • the turning vane arrangement 30 employs a forced or pressurized flow to reduce engine backpressure, and improve engine performance.
  • the integration of the turning vane arrangement 30 with at least one other the flow source 54 aboard the vehicle, i.e., the integrated system, improves the overall IR and radar detectability of the aircraft. That is, the fluid communication means 54. 56 provides a structure which occludes the orifices which would otherwise be sources of IR energy or radar reflectivity.
  • nozzle configuration 60 R While the exhaust nozzle has been described as having a refective lobe configuration 60 R , other nozzle configurations which are useful include an aligned lobe configuration 60 A such as depicted in Fig. 5a, and a penetrating lobe configuration 60 P such as shown in Fig. 5b. Similar to the reflective lobe configuration 60 R , these other lobe configurations 60 A , 60 P utilize the same mixing principle.
  • the described embodiment includes a dual function shroud/duct 32, i.e., a shroud for producing a first stage ejector and a duct for directing flow downstream
  • the turning vane arrangement 30 may be used in combination with a simple duct, i.e., a duct that does not also function as a shroud.
  • the described embodiment shows the exterior conduit 56 of the fluid communication means 54, 56 as being connected to an upper side of the interior conduit, it will be appreciated that the exterior conduit may be bifurcated to feed both sides, i.e., upper and lower sides, of the interior conduit 54.
  • While the described embodiment describes the integration of the turning vane arrangement 30 with the pressurized flow of a transmission oil cooler, another source or multiple sources of pressurized air or exhaust may used.
  • engine bleed air or air conditioning exhaust may be of sufficiently low temperature to ameliorate the temperature characteristics of the primary exhaust flow.
  • it may be desirable to integrate multiple other sources into a single system to reduce both IR and radar detectability.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Exhaust Gas After Treatment (AREA)

Abstract

Cet agencement de déflecteur (30), destiné à un dispositif de suppression d'infrarouges présentant un conduit (32) servant à recevoir un premier écoulement (PF) de gaz d'échappement du moteur, fonctionne pour diriger un second écoulement sous pression (PSF) de fluide gazeux à basse température, dans le premier écoulement de gaz d'échappement, aux fins de réduction de l'énergie infrarouge de ceux-ci. Cet agencement (30) comprend au moins un déflecteur (50) placé dans le conduit (32) et caractérisé par un bord de fuite (50TE) délimitant une tuyère d'éjection des gaz (52). Cette tuyère (52) est conçue pour être en communication fluidique avec une source d'écoulement (54) produisant le second flux sous pression, et elle peut comporter plusieurs lobes contigus (60R, 60A ou 60P) destinés à mélanger rapidement et complètement le premier écoulement (PF) et le second écoulement sous pression (PSF).
PCT/US1998/014212 1997-07-23 1998-07-08 Agencement de deflecteur pour dispositif de suppression d'infrarouges WO1999005407A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
DE69805800T DE69805800T2 (de) 1997-07-23 1998-07-08 Umlenkschaufel für eine einrichtung zur infrarotstrahlungsabschirmung
AU82957/98A AU8295798A (en) 1997-07-23 1998-07-08 Turning vane arrangement for ir suppressors
JP2000504365A JP2001511495A (ja) 1997-07-23 1998-07-08 赤外線抑制のための転向ベーン配列
EP98933276A EP0998628B1 (fr) 1997-07-23 1998-07-08 Agencement de deflecteur pour dispositif de suppression d'infrarouges

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US08/898,904 1997-07-23
US08/898,904 US6055804A (en) 1997-07-23 1997-07-23 Turning vane arrangement for IR suppressors

Publications (1)

Publication Number Publication Date
WO1999005407A1 true WO1999005407A1 (fr) 1999-02-04

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Application Number Title Priority Date Filing Date
PCT/US1998/014212 WO1999005407A1 (fr) 1997-07-23 1998-07-08 Agencement de deflecteur pour dispositif de suppression d'infrarouges

Country Status (6)

Country Link
US (1) US6055804A (fr)
EP (1) EP0998628B1 (fr)
JP (1) JP2001511495A (fr)
AU (1) AU8295798A (fr)
DE (1) DE69805800T2 (fr)
WO (1) WO1999005407A1 (fr)

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EP1750003A3 (fr) * 2005-08-03 2010-09-08 General Electric Company Dispositif permettant la supression d'infra-rouges
EP1750002A3 (fr) * 2005-08-03 2010-09-08 General Electric Company Dispositif compenant un déflecteur mono-pièce permettant la suppression d' infra-rouges
CN104033283A (zh) * 2014-06-05 2014-09-10 中国航空工业集团公司沈阳发动机设计研究所 一种新型二元单边膨胀喷管
EP3362660A4 (fr) * 2015-10-13 2019-06-12 Sikorsky Aircraft Corporation Dispositif pour modifier une signature infrarouge d'un conduit d'échappement et procédé pour modifier une signature infrarouge d'un conduit d'échappement

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JP5013710B2 (ja) * 2005-12-22 2012-08-29 ゼネラル・エレクトリック・カンパニイ 赤外線抑制装置及び方法
US7836682B2 (en) * 2006-02-13 2010-11-23 General Electric Company Methods and apparatus for operating a pulse detonation engine
WO2007106188A2 (fr) * 2006-02-27 2007-09-20 Sikorsky Aircraft Corporation Système de suppression de la signature infrarouge avec cloison spirale
US20100199626A1 (en) * 2008-12-31 2010-08-12 Benjamin Roland Harding Turbine engine exhaust gas tube mixer
US20100326049A1 (en) * 2009-06-25 2010-12-30 Honeywell International Inc. Cooling systems for rotorcraft engines
RU2490496C2 (ru) * 2011-11-10 2013-08-20 Открытое акционерное общество "Научно-производственное объединение "Сатурн" (ОАО "НПО "Сатурн") Выходное устройство двухконтурного газотурбинного двигателя
EP2719620A3 (fr) 2012-10-10 2017-10-18 Sikorsky Aircraft Corporation Système d'échappement tourné vers le haut pour aéronef à voilure tournante
US9051047B2 (en) * 2012-11-12 2015-06-09 Sikorsky Aircraft Corporation Flexbeam rotor attachment to rotor blade
US9593691B2 (en) 2013-07-19 2017-03-14 General Electric Company Systems and methods for directing a flow within a shroud cavity of a compressor
CN110748436B (zh) * 2019-10-16 2021-06-15 南京航空航天大学 一种双层冷却混合管引射式红外抑制器
US11608189B2 (en) * 2020-08-11 2023-03-21 Textron Innovations Inc Multistage infrared suppression exhaust system
CN112065604B (zh) * 2020-08-31 2021-11-30 南京航空航天大学 一种低红外特征喷管
US20220154613A1 (en) * 2020-11-17 2022-05-19 Pacific Crown Helicopters Deflector exhaust nozzle for as350/ec130 and for other single engine helicopters
US11434018B1 (en) * 2021-11-27 2022-09-06 Airbus Defence and Space GmbH Switchable air inlet device for engine air

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EP0998628A1 (fr) 2000-05-10
EP0998628B1 (fr) 2002-06-05
JP2001511495A (ja) 2001-08-14
US6055804A (en) 2000-05-02
DE69805800D1 (de) 2002-07-11
DE69805800T2 (de) 2003-02-20
AU8295798A (en) 1999-02-16

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